Compressor for a gas turbine engine

ABSTRACT

A compressor for a gas turbine engine comprising a casing, a mixed flow stage including a first rotor and at least one further mixed or centrifugal flow stage including a second rotor axially spaced apart from the first rotor, and bearing means positioned between the first and second rotors for rotatably mounting the rotors with respect to the casing. The bearing means is positioned closely adjacent the first rotor. The rotors are carried on a common shaft and the bearing means rotatably mounts the shaft with respect to a bearing support which extends radially with respect to the shaft between the bearing and the casing. The bearing support comprises a frusta-conical portion extending from the said bearing means towards the said casing.

This invention relates to a compressor for a gas turbine engine

In particular a two stage compressor with a mixed flow first stage andat least one further stage for a single spool gas turbine engine.

Compressors for gas turbine engines include axial flow and radial flowtypes both of which have their particular advantages depending upon thespecific engine application. For example, it is possible to reduce thenumber of compressor stages, and hence cost, by the use of a singlecentrifugal flow stage instead of a greater number of axial flow stages.Stage for stage much higher pressure ratios are possible withcentrifugal compressors than axial flow machines and centrifugal flowcompressors have greater resistance to foreign object damage byingestion of objects in the intake air stream. Centrifugal flowcompressors are more commonly found in small turboshaft type engineswhere resistance to such damage can be a highly relevant designconsideration, for example in helicopter applications. Axial flowcompressors on the other hand offer greater operational efficiency atthe expense of more compressor stages, and hence components and cost,for a given pressure ratio. Axial flow compressors also have a lowerfrontal area than centrifugal flow compressors which can be an importantconsideration in aircraft applications. The increased frontal area, orenvelope, associated with radial flow compressors is due in part to thedimensions of the radial flow impeller and also the requirement toposition an annular diffuser radially outwards of the impeller.

The conflicting requirements of compressor operational efficiency andreduced number of compressor stages has been partly addressed byso-called mixed flow stages in which the flow through the compressorstage has both an axial and radial component so that the stage functionspartly as a radial flow stage and partly as an axial flow stage. Mixedflow compressors offer a combination of the performance benefits ofaxial flow and radial flow compressors.

A two-stage compressor with a mixed flow first stage and a centrifugalflow second stage is disclosed in International Patent ApplicationNumber PCT/CA01/01335, in which the mixed flow rotor and centrifugalrotor are joined together on a common shaft which is supported withrespect to the housing or compressor casing by means of a bearingassembly positioned at the forward (upstream end) of the two stagecompressor. In this arrangement the compressor rotor stages may beconsidered to be cantilevered from the bearing at the forward end of thecompressor and this can lead to problems in particular with rotor bladetip clearances. This arrangement also has the disadvantage that itrequires a bearing support structure to be provided in the inlet regionof the compressor resulting in a reduced flow cross-section, or agreater diameter inlet, or compressor frontal area, for a specific flowcross sectional area.

There is a requirement therefore for a compressor having at least afirst mixed flow stage and a second mixed or centrifugal flow stagewhich exhibits improved tip clearance characteristics, and also arequirement for such a compressor having a reduced diameter inlet for aspecific flow cross sectional area.

According to an aspect of the invention there is provided a compressorfor a gas turbine engine; the compressor comprising a casing, a bearingmeans, a bearing support means, a mixed flow stage including a firstrotor and at least one further mixed or centrifugal flow stage includinga second rotor axially spaced apart from the said first rotor; all ofthe said rotors being carried on a common shaft rotatably mounted by thebearing means with respect to the bearing support means, the saidbearing means being positioned between said first and second rotors forrotatably mounting the said rotors with respect to the said casing.

This arrangement has the particular advantage that the bearing and itsassociated support structure can be positioned away from the compressorinlet, at the rear face, that is to say the downstream side, of thefirst rotor so that the axial distance between the bearing and the rotorblades, or vanes, is significantly less than in arrangements where therotors are effectively cantilevered from the front of the compressor aspreviously described. In this way it is readily possible to maintain anappropriate tip clearance between the compressor rotor vanes and thecompressor casing or other ducting forming the radially outer wall ofthe annular gas flow passage of the compressor. This arrangement has theadditional benefit of reducing the compressor inlet diameter for aspecific inlet flow area since the radially inner flow boundary of theinlet may be positioned closer to the rotation axis of the compressor.In this aspect of the invention the bearing and associated supportstructure is positioned remote from the compressor inlet and placedbetween the first and second stage rotors, that is to say between thefront face of the first stage rotor and the rear face of the secondstage rotor.

By positioning the bearing between the first and second stage rotors thebearing support can be positioned downstream of the first compressorstage well away from the inlet section where it would reduce the flowinlet area for a particular diameter of inlet.

In preferred embodiments the bearing means is positioned closelyadjacent to the first rotor. This arrangement has the additional benefitof readily enabling a frusto-conical bearing support structure to bepositioned between the first and second stage rotors to support thebearing loads with respect to the compressor casing. This arrangementhas particular advantages where the second stage comprises a centrifugalflow compressor stage where radial tip clearances are of lesssignificance than in the first stage mixed flow impeller.

The bearing support means preferably extends radially with respect tothe compressor shaft between the bearing and the casing so that thebearing loads can be readily transferred to the compressor casing, andpreferably the bearing support is integrated, that is in the sense thatit forms part of rather than being integrally formed with, with othernon rotating components within the gas flow path of the compressorbetween the first and second stage rotors.

In one embodiment the bearing support means comprises a frusto-conicalportion extending from the bearing towards the casing. This isparticularly advantageous when the bearing support transfers axialthrust loads from the compressor shaft to the casing. In particularembodiments an angle of 45° may be an appropriate angle for the conicalportion.

Preferred embodiments comprise at least one diffuser section between thefirst and second stage rotors. In arrangements where such a diffusersection is provided the bearing support may be connected to thecompressor casing through the diffuser section, that is to say throughstructural vane support struts extending radially between the inner andouter radial peripheries of the annular gas flow path of the compressor.

Preferably, the diffuser section has an inlet and an outlet with aradially inward flow section therebetween. In this way it is possible toredirect the flow at the exit of the first stage radially inwards sothat the inlet to the second stage rotor is at a position radiallyinward of the first stage diffuser outlet. This readily enables thediameter of the compressor to be reduced in the region of the secondstage at least, so that the frontal area of the compressor, asdetermined by the compressor casing, can be optimised, ie kept to aminimum.

In one particular embodiment the bearing comprises a journal bearing andin other embodiments additionally or alternatively a thrust bearing.

In one embodiment the compressor is a two staged mixed flow compressorhaving two mixed flow stages. In such an embodiment the compressorstages may have substantially equal pressure ratios preferably greaterthan 4:1 so that the two stage compressor has an overall pressure ratioof at least 16:1.

According to another aspect of the invention there is provided a gasturbine engine comprising a compressor having a mixed flow stage and atleast one further mixed or centrifugal stage according to the compressoras hereinbefore defined in accordance with the above first mentionedaspect of the invention.

In one preferred embodiment the gas turbine engine is a single spoolengine, for example a turbofan or turbojet single spool engine. For theavoidance of doubt the term “single spool” used herein refers to anengine configuration in which all the turbine and compressor rotorstages are mounted on a common engine shaft. Such an engineconfiguration is particularly suitable for applications where a low costengine is required, for example where production and maintenance costsare a more significant consideration than operating costs, for examplewhere the engine has a limited operational life. The gas turbine enginemay be an aero engine for a manned or unmanned aircraft as required.

An embodiment of the invention will now be more particularly described,by way of example, with reference to the accompanying drawing, whichshows an axial cross-section view through part of a single spool gasturbine engine comprising a compressor according to one embodiment ofthe present invention.

Referring to the drawing a single spool gas turbine engine comprises atwo stage mixed flow compressor 10 including a first stage 12 and asecond stage 14 mounted within a generally cylindrical compressorhousing 16. The compressor 10 comprises part of a single spool gasturbine engine, the other components of which are not shown in thedrawing but briefly comprise a combustor downstream of the compressorand a turbine downstream of the combustor and rotatably connected to thecompressor rotor stages by a common shaft 18. The shaft 18 is rotatablymounted with respect to the compressor housing or casing 16 by a firstbearing assembly 20 located in the compressor, and additionally a secondbearing assembly (not shown) downstream of the compressor, for examplein the turbine section of the engine. The compressor 10 has an inlet 22through which air is inducted into the first stage 12, and an outlet 24through which high pressure air is delivered to the engine combustorsection (not shown) downstream of the compressor.

The first compressor stage 12 comprises a rotor 26 which is connected tothe forward end of the shaft 18 and a diffuser 28. The rotor 26 includesa disc part 30 which carries a plurality of circumferentially spacedvanes 32 which pressurise the inducted air in the annular gas flowpassage defined between the rim 34 of the disc and an interior wall part36 of the gas flow passage. The mixed flow rotor turns the inductedinlet air so that the air which has a generally axial flow direction atthe inlet has an outlet flow direction which includes a radial as wellas an axial component. As it is well understood in the art of gasturbine engine design the diffuser 28 functions to reduce the outletvelocity of the gas flow exiting the rotor vanes 32 so that the flow maybe delivered at an appropriate velocity to a downstream compressorstage. In the embodiment shown in the drawing the diffuser 28 turns thegas flow exiting the rotor vanes 32 so that it exits the diffuser at 38having a substantially axial flow direction.

The second mixed flow compressor stage 14 is substantially the same asthe first compressor stage in the sense that it comprises a rotor 42including a disc 44 and vanes 46 and a diffuser section 48 immediatelydownstream of the rotor vanes 46. The two mixed flow stages 12 and 14are spaced apart along the compressor axis 50 with the outlet 38 to thefirst stage diffuser 28 being connected to the second stage rotor inlet52 by a further duct 40 which extends between the two compressor stages,and to a bypass duct 66 radially outwards of the duct 40. The inner andouter annular surfaces of the duct 40 comprise inflection curves ofrevolution such that the duct 40 turns the flow radially inwards fromthe diffuser outlet 38 to the second stage inlet 52 which has a smallermean radius with respect to the compressor arcs 50 than the outlet 38.The duct 40 preferably comprises an array of support vanes 54 whichextend radially through the gas flow passage of the duct to provide astructural support between the casing 16 and a bearing support assembly56 which extends from the vane structure 54 to the bearing assembly 20.The support vanes 54 may extend only partially along the duct 40 asshown in the drawing, or alternatively extend along the full length ofthe duct between first stage outlet 38 and second stage inlet 52. In theembodiment shown in the drawing the vanes 54 are connected to thecompressor casing through the bypass duct 66 by corresponding radiallyextending bypass support vanes 62.

The bearing support 56 includes a frusto-conical part 58 and acylindrical part 60 and two radial parts 62 and 64. The annularstructure defined by the bearing support parts 58, 60, 62 and 64 readilysupports the shaft and rotor stages connected thereto and provides forthe transfer of engine loads from the rotatable components to thecompressor casing 16. The radial parts 62 and 64 schematicallyillustrate the axial extent of the vanes 54 when they extend the fulllength of the duct 40.

In the drawing the bearing assembly 20 is positioned between thecompressor rotors 26 and 42 and in this particular embodiment it ispositioned towards the rear face, that is to say the downstream side, ofthe first stage rotor 26. In other embodiments the bearing assembly 20may be positioned anywhere between the rotors 26 and 42 withoutsignificantly affecting the tip clearance control of the rotor vanes 32with respect to the outer wall 36 of the annular gas flow passage of thecompressor.

Although aspects of the invention have been described with reference tothe embodiments shown in the accompanying drawing, it is to beunderstood that the invention is not limited to this precise embodimentand that various changes and modifications may be effected withoutfurther inventive skill and effort. For example, the second compressorstage 14 may comprise a centrifugal flow compressor instead of a secondmixed flow compressor. In addition the gas turbine engine may comprise aturbofan or turbojet engine. In alternative embodiments the bearingsupport structure 56 may be supported by the compressor casing 16through the diffuser section 28 rather than the duct 40. Moreover, oneor more additional bearing assemblies may be provided between thecompressor rotors 26 and 42 if additional support is required. Theinvention also contemplates embodiments where further mixed flow orcentrifugal flow compressor stages are provided, that is to say three,four, or more compressor stages.

1. A compressor for a gas turbine engine; the compressor comprising acasing, a bearing means, a bearing support means, a mixed flow stageincluding a first rotor and at least one further mixed or centrifugalflow stage including a second rotor axially spaced apart from the saidfirst rotor; all of the said rotors being carried on a common shaftrotatably mounted by the bearing means with respect to the bearingsupport means, the said bearing means being positioned between saidfirst and second rotors for rotatably mounting the said rotors withrespect to the said casing.
 2. A compressor as claimed in claim 1wherein the said bearing means is positioned closely adjacent the saidfirst rotor.
 3. A compressor as claimed in claim 1 wherein the saidbearing support means extends radially with respect to the said shaftbetween the said bearing means and the said casing.
 4. A compressor asclaimed in claim 3 wherein the said bearing support means comprises afrusta-conical portion extending from the said bearing means towards thesaid casing.
 5. A compressor as claimed in claim 1 further comprising atleast one diffuser section between the said rotors and wherein the saidbearing support means is connected to the said casing through the saiddiffuser section.
 6. A compressor as claimed in claim 5 wherein the saiddiffuser section has an inlet section and an outlet section with aradially inward flow section there between.
 7. A compressor as claimedin claim 1 wherein the bearing means comprises a journal bearing.
 8. Acompressor as claimed in claim 7 wherein the bearing means additionallycomprises a thrust bearing.
 9. A compressor as claimed in claim 1wherein the said compressor is a two stage mixed flow compressor havingtwo mixed flow stages.
 10. A compressor as claimed in claim 1 whereinthe said compressor stages have substantially the same pressure ratio.11. A compressor as claimed in claim 10 wherein the pressure ratio ofeach stage is greater than 4:1.
 12. A gas turbine engine comprising acompressor according to claim
 1. 13. A gas turbine engine as claimed inclaim 12 wherein the said origin is a single spool engine.
 14. A gasturbine engine as claimed in claim 13 wherein the said engine comprisesa turbo-fan or a turbo-jet engine.
 15. A gas turbine engine as claimedin claim 12 wherein the said engine is for an aircraft.
 16. A gasturbine engine as claimed in claim 13 wherein the said engine is for anaircraft.
 17. A gas turbine engine as claimed in claim 14 wherein thesaid engine is for an aircraft.